U.S. patent application number 16/221910 was filed with the patent office on 2022-09-22 for damping device.
This patent application is currently assigned to Safran Aircraft Engines. The applicant listed for this patent is Safran Aircraft Engines. Invention is credited to Francois Jean COMIN, Charles Jean-Pierre DOUGUET, Laurent JABLONSKI, Philippe Gerard Edmond JOLY, Romain Nicolas LAGARDE, Marie-Oceane PARENT, Jean-Marc Claude PERROLLAZ.
Application Number | 20220299040 16/221910 |
Document ID | / |
Family ID | 1000005751425 |
Filed Date | 2022-09-22 |
United States Patent
Application |
20220299040 |
Kind Code |
A1 |
JOLY; Philippe Gerard Edmond ;
et al. |
September 22, 2022 |
DAMPING DEVICE
Abstract
A turbomachine assembly is provided. The assembly includes: a
first rotor module including a first blade, a second rotor module,
connected to the first rotor module, and including a second blade
of smaller length than the first blade, and a damping device
including a plurality of first radial external surfaces supported
with friction against the first module. The damping device is
stepped, and includes a second radial external surface supported
with friction against the second module, so as to couple the
modules for the purpose of damping their respective vibrational
movements during operation.
Inventors: |
JOLY; Philippe Gerard Edmond;
(Moissy-Cramayel, FR) ; PARENT; Marie-Oceane;
(Moissy-Cramayel, FR) ; COMIN; Francois Jean;
(Moissy-Cramayel, FR) ; LAGARDE; Romain Nicolas;
(Moissy-Cramayel, FR) ; PERROLLAZ; Jean-Marc Claude;
(Moissy-Cramayel, FR) ; JABLONSKI; Laurent;
(Moissy-Cramayel, FR) ; DOUGUET; Charles Jean-Pierre;
(Moissy-Cramayel, FR) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Safran Aircraft Engines |
Paris |
|
FR |
|
|
Assignee: |
Safran Aircraft Engines
Paris
FR
|
Family ID: |
1000005751425 |
Appl. No.: |
16/221910 |
Filed: |
December 17, 2018 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F04D 29/34 20130101;
F01D 5/10 20130101; F05D 2220/36 20130101; F05D 2260/96 20130101;
F01D 5/22 20130101; F01D 5/26 20130101; F04D 29/322 20130101 |
International
Class: |
F04D 29/34 20060101
F04D029/34; F04D 29/32 20060101 F04D029/32 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 20, 2017 |
FR |
17 62586 |
Claims
1. A turbomachine assembly comprising: a first rotor module
comprising a first blade, a second rotor module, connected to the
first rotor module at a connection, and comprising a second blade
of smaller length than the first blade, and a damping device
comprising a plurality of first radial external surfaces supported
with friction against the first rotor module, wherein the damping
device is stepped and comprises a second radial external surface
supported with friction against the second rotor module, so as to
couple the the first rotor module with the second rotor module in
order to dampen a vibrational movement of the first rotor module
with respect to the second rotor module during operation, and
wherein a bore is formed both into a farthest downstream first
radial external surface among the plurality of first radial
external surfaces, and into the second radial external surface so
as to form a U-shaped surface.
2. The turbomachine assembly of claim 1, wherein the damping device
further comprises an abutment surface supported against the
connection between the first rotor module and the second rotor
module, so as to accomplish an axial retention of the damping
device.
3. The turbomachine assembly of claim 1, wherein: the first rotor
module comprises a disk being centered on a turbomachine
longitudinal axis and presenting an external periphery, the first
blade being mounted on the external periphery of the disk from
which if the first blade extends, the first blade further
comprising an airfoil, a platform, a support and a root embedded in
a recess of the disk, and the second rotor module comprises a
ferrule comprising a circumferential extension extending toward the
platform of the first blade, the plurality of first radial external
surfaces of the damping device being supported with friction on a
plurality of respective internal surfaces of the platform of the
first blade, the second radial external surface of the damping
device being supported with friction against the circumferential
extension of the ferrule of the second rotor module.
4. (canceled)
5. The turhomachine assembly of claim 1, wherein each first radial
external surface as well as the second radial external surface of
the damping device are formed respectively from a sacrificial plate
configured to guarantee a support with friction of the plurality of
first radial external surfaces and the second radial external
surfaces.
6. The turbomachine assembly of claim 2, wherein the abutment
surface is formed from a sacrificial plate configured to guarantee
an axial support of the abutment surface.
7. The turbomachine assembly of claim 5, wherein the sacrificial
plates comprise a dissipative coating.
8. The turbomachine assembly according of claim 5, wherein the
sacrificial plates comprise a viscoelastic coating.
9. The turbomachine assembly according of claim 5, wherein the
first rotor module comprises a disk being centered on a
turbomachine longitudinal axis, and wherein the sacrificial plate
forming the farthest downstream radial external surface has a
variable thickness along the turbomachine longitudinal axis.
10. The turbomachine assembly of claim 1, wherein the damping
device comprises lightening bores designed to lighten the damping
device.
11. The turbomachine assembly of claim 1, wherein the damping
device comprises inserts, designed to add weight to the damping
device.
12. The turbomachine assembly of claim 1, wherein the first module
is a fan, and the second module is a low-pressure compressor.
13-14. (canceled)
Description
TECHNICAL FIELD
[0001] The invention relates to an assembly comprising a
turbomachine rotor module.
[0002] The invention relates more specifically to an assembly for a
turbomachine comprising two rotor modules and a damping device.
PRIOR ART
[0003] A turbomachine rotor module generally comprises one or more
stage(s), each stage comprising a disk centered on a turbomachine
longitudinal axis, corresponding to the axis of rotation of the
rotor module. The rotation of the disk is generally ensured by a
rotating shaft to which it is integrally connected, for example by
means of a rotor module trunnion, the rotating shaft extending
along the turbomachine longitudinal axis. Blades are mounted on the
external periphery of the disk, and distributed circumferentially
in a regular manner around the longitudinal axis. Each blade
extends from the disk, and also comprises an airfoil, a platform, a
support and a root. The root is embedded in a recess of the disk
configured for this purpose, the airfoil is swept by a flow passing
through the turbomachine and the platform forms a portion of the
internal surface of the flow path.
[0004] The range of operation of a rotor module is limited, in
particular due to aeroelastic phenomena. The rotor modules of
modern turbomachines, which have a high aerodynamic loading and a
reduced number of blades, are more sensitive to this type of
phenomena. In particular, they have reduced margins between the
operating zones without instability and the unstable zones. It is
nevertheless imperative to guarantee a sufficient margin between
the stability range and that of instability, or to demonstrate that
the rotor module can operate in the unstable zone without exceeding
its endurance limit. This allows guaranteeing risk-free operation
over its entire life and the entire range of operation of the
turbomachine.
[0005] Operation in the zone of instability is characterized by
coupling between the fluid and the structure, the fluid applying
the energy to the structure, and the structure responding with its
natural modes at levels which can exceed the endurance limit of the
material constituting the blade. This generates vibrational
instabilities which accelerate the wear of the rotor module and
reduce its lifetime.
[0006] In order to limit these phenomena, it is known to implement
a system damping the dynamic response of the blade, so as to
guarantee that it does not exceed the endurance limit of the
material, regardless of the operating point of the rotor module.
However, most of the known systems of the prior art are dedicated
to dampen vibration modes with non-zero dephasing, and
characterizing an asynchronous response of the blades to
aerodynamic forces. Such systems have for example been described in
documents FR 2 949 142, EP 1 985 810 and FR 2 923 557, in the
Applicant's name. These systems are all configured to be
accommodated between the platform and the root of each blade, in
the recess delimited by the respective supports of two successive
blades. Moreover, such systems operate, when two successive blade
platforms are moved with respect to one another, by dissipating the
vibration energy, by friction for example.
[0007] However, these systems are completely ineffective for
damping vibration modes having a zero-dephasing involving the
blades and the rotor line, i.e. its rotating shaft. Such modes are
characterized by a flexure of the rotor blades with zero
inter-blade dephasing implying a non-zero moment on the rotating
shaft. In addition, this is a coupled mode between the blade, the
disk and the rotating shaft. More precisely, the torsion within the
rotor module, resulting for example from reverse forces between a
turbine rotor and a compressor rotor, lead to flexural movements of
the blades with respect to their attachment to the disk. These
movements are greater the longer the blade, and the more the
attachment is flexible.
[0008] Thus, there exists a need for a damping system for a
turbomachine rotor making it possible to limit the instabilities
generated by all modes of vibration as previously described.
SUMMARY OF THE INVENTION
[0009] One goal of the invention is to dampen vibration modes with
zero dephasing for all types of turbomachine rotor modules.
[0010] Another goal of the invention is to influence the damping of
vibration modes with non-zero dephasing, for all types of
turbomachine rotor modules.
[0011] Another goal of the invention is to propose a damping
solution that is simple and easy to implement.
The invention proposes in particular a turbomachine assembly
comprising: [0012] a first rotor module comprising a first blade,
[0013] a second rotor module, connected to the first rotor module,
and comprising [0014] a second blade of smaller length than the
first blade, and [0015] a damping device comprising a plurality of
first radial external surfaces supported with friction against the
first module, characterized in that the damping device is stepped,
and comprises a second radial external surface supported with
friction against the second module, so as to couple the modules for
the purpose of damping their respective vibrational movements
during operation.
[0016] The mechanical coupling between the first and the second
rotor module allows increasing the tangential stiffness of the
connection between these two rotors, while still allowing a certain
axial and radial flexibility of the damping device so as to
maximize contact between the different elements of the assembly.
This makes it possible to limit the instabilities related to the
vibration mode with zero dephasing, but also to participate in the
damping of vibration modes with non-zero dephasing. In addition,
such an assembly has the advantage of an easy integration within
existing turbomachines, whether during manufacture or during
maintenance. Indeed, the stepped structure of the damping device
allows easier assembly, for example at the internal surface of the
platform of a fan blade.
[0017] The assembly according to the invention can also comprise
the following features, taken alone or in combination: [0018] the
damping device also comprises an abutment surface supported against
the connection between the first and the second module, so as to
accomplish the axial retention of the damping device, [0019] the
first rotor module comprises a disk centered on a turbomachine
longitudinal axis, the first blade being mounted on the radial
external periphery of the disk from which it extends, and also
comprising an airfoil, a platform, a support and a root embedded in
the recess of the disk, and the second module comprises a ferrule
comprising a circumferential extension extending toward the
platform of the first blade, the plurality of first radial external
surfaces of the damping device being supported with friction on a
plurality of respective internal surfaces of the platform of the
first blade, the second radial external surface of the damping
device being supported with friction against the circumferential
extension of the ferrule of the second rotor module, [0020] the
farthest downstream first radial external surface among the
plurality of first radial external surfaces, and the second radial
external surface, are dug by a bore, so as to form a U-shaped
surface, [0021] each first radial external surface as well as the
second radial external surface of the damping device are formed
respectively from a sacrificial plate configured to guarantee the
support with friction of said surfaces, [0022] the abutment surface
is formed from a sacrificial plate configured to guarantee the
axial support of the abutment surface, [0023] the plates comprise a
coating of the dissipative type, [0024] the plates comprise a
coating of the viscoelastic type, [0025] the plate forming the
farthest downstream radial external surface has a variable
thickness along the turbomachine longitudinal axis, [0026] the
damping device comprises bores designed to lighten the damping
device, [0027] the damping device comprises inserts, of the
metallic type for example, designed add weight to the damping
device, and [0028] the first module is a fan, and the second module
a compressor, for example a low-pressure compressor.
[0029] The invention also relates to a turbomachine comprising an
assembly as previously described.
[0030] The invention also relates to a stepped damping device
comprising a plurality of first radial external surfaces configured
to be supported with friction against a first module of an assembly
as previously described, and also comprising a second radial
external surface configured to be supported with friction against a
second module of such an assembly, so as to couple the modules for
the purpose of damping their respective vibrational movements
during operation.
RAPID DESCRIPTION OF THE FIGURES
[0031] Other features, goals and advantages of the present
invention will appear upon reading the detailed description that
follows and with reference to the appended drawings given by way of
non-limiting examples and in which:
[0032] FIG. 1 is a schematic section view of an exemplary
embodiment of the assembly according to the invention,
[0033] FIG. 2 is a front view of a rotor module subjected to
tangential vibrations of which the mode has zero dephasing,
[0034] FIG. 3a illustrates schematically tangential movements of
the turbomachine rotor modules, as a function of the position of
said modules along a turbomachine axis,
[0035] FIG. 3b is an enlargement in schematic perspective of the
interface between two turbomachine rotor modules illustrating its
tangential movements relative to said rotor modules,
[0036] FIG. 4a illustrates schematically a first exemplary
embodiment of a damping device according to the invention,
[0037] FIG. 4b illustrates schematically a second exemplary
embodiment of a damping device according to the invention,
[0038] FIG. 5a illustrates schematically a third exemplary
embodiment of a damping device according to the invention, and
[0039] FIG. 5b illustrates schematically a fourth exemplary
embodiment of a damping device according to the invention.
DETAILED DESCRIPTION OF THE INVENTION
[0040] An exemplary embodiment of an assembly 1 according to the
invention will now be described, with reference to the figures.
[0041] Hereafter, upstream and downstream are defined with respect
to the normal flow direction of air through the turbomachine.
Furthermore, a turbomachine longitudinal axis X-X is defined. In
this manner, the axial direction corresponds to the direction of
the turbomachine longitudinal axis X-X, a radial direction is a
direction which is perpendicular to this turbomachine longitudinal
axis X-X and which passes through said turbomachine longitudinal
axis X-X, and a circumferential direction corresponds to the
direction of a closed planar curve, of which all points are located
at equal distance from the turbomachine longitudinal axis X-X.
Finally, and unless the contrary is stated, the terms "internal (or
interior)" and "external (or exterior)" respectively, are used with
reference to a radial direction so that the internal (i.e. radially
internal) portion or face of an element is closer to the
turbomachine longitudinal axis X-X than the external (i.e. radially
external) portion or face of the same element.
[0042] Referring to FIGS. 1 and 3a, such an assembly 1 comprises:
[0043] a first rotor module 2 comprising a first blade 20, [0044] a
second rotor module 3, connected to the first rotor module 2, and
comprising a second blade 30 with a length smaller than the first
blade 20, and [0045] a stepped damping device 4 comprising: [0046]
a plurality of first radial external surfaces 40, 42, 44 supported
with friction against the first module 2, and [0047] a second
radial external surface 41 supported with friction against the
second module 3, [0048] so as to couple the modules for the purpose
of damping their respective vibrational movements during
operation.
[0049] By support "with friction" is meant that the contact between
the radial external surfaces 41, 42 and, respectively, the first
rotor module 2 and the second rotor module 3 occurs with friction.
In other words, the support forces between the radial external
surfaces 41, 42 and, respectively, the first rotor module 2 and the
second rotor module 3 can be decomposed into pressure forces which
are directed normal to the contact, and friction forces, directed
tangentially to the contact. This support guarantees both the
mechanical consistency of the assembly 1, by means of the pressure
forces, but also the coupling between the modules 2, 3 for the
purpose of damping their respective vibrational movements during
operation, by means of the friction forces.
[0050] Referring to FIGS. 1 and 3a, the first rotor module is a fan
2, and the second rotor module is a low-pressure compressor 3,
situated immediately downstream of the fan 2.
[0051] The fan 2 and the low-pressure compressor 3 comprise a disk
21, 31 centered on the turbomachine longitudinal axis X-X, the
first 20 and the second 30 blade being respectively mounted on the
external periphery of the disk 21, 31 and also comprising an
airfoil 23, 33, a platform 25, 35, a support 27, 37 and a root 29,
39 embedded in a recess 210, 310 of the disk 21, 31. The distance
separating the root 29, 39 from the end of the airfoil 23, 33
constitutes the respective lengths of the first 20 and of the
second 30 blade. The length of the first blade 20 and second blade
30 is therefore considered here to be substantially radial with
respect to the longitudinal axis X-X of rotation of the rotor
modules 2, 3. In operation, the blade 23, 33 is swept by a flow 5
passing through the turbomachine, and the platform 25, 35 forms a
portion of the internal surface of the flow path 5. Generally, as
can be seen in FIGS. 2 and 3a, the fan 2 and the low-pressure
compressor 3 comprise a plurality of blades 20, 30 distributed
circumferentially around the longitudinal axis X-X. The
low-pressure compressor 3 also comprises an annular ferrule 32 also
centered on the longitudinal axis X-X. The ferrule 32 comprises a
circumferential extension 34, also annular, extending toward the
platform 25 of the first blade 20. This annular extension 34
carries radial knife edge seals 36 configured to prevent air flow
rate losses from the flow path 5. Moreover, the ferrule 32 is
attached to the disk 21 of the fan 2 by means of attachments 22
distributed circumferentially around the longitudinal axis X-X.
Such attachments can for example be bolted connections 22.
Alternatively, such attachments 22 can be achieved by an
interference fit to which is associated an anti-rotation device
and/or an axial locking system. Finally, with reference to FIG. 3a,
the assembly formed from the fan 2 and the compressor 3 is rotated
by a rotating shaft 6, called the low-pressure shaft, to which the
fan 2 and the low-pressure compressor 3 are integrally connected,
by means of a rotor trunnion 60, the low-pressure shaft 6 being
also connected to a low-pressure turbine 7, downstream of the
turbomachine, and extending along the turbomachine longitudinal
axis X-X.
[0052] In operation, the fan 2 aspires air of which all or part is
compressed by the low-pressure compressor 3. The compressed air
then circulates in a high-pressure compressor (not shown) before
being mixed with fuel, then ignited within the combustion chamber
(not shown), to finally be successively expanded in the
high-pressure turbine (not shown), and the low-pressure turbine 7.
The opposite forces of compression, upstream and of expansion
downstream cause aeroelastic flutter phenomena, which couple the
aerodynamic forces on the blades 20, 30 and the flexural and
torsional vibration movements in the blades 20, 30. As illustrated
in FIG. 2, this flutter causes in particular intense torsional
forces within the low-pressure shaft 6 which are fed through to the
fan 2 and to the low-pressure compressor 3. The blades 20, 30 are
then subjected to tangential pulses, particularly according to a
vibration mode with zero dephasing. This is in fact a flexural mode
with zero inter-blade 20, 30 dephasing, involving a non-zero moment
on the low-pressure shaft 6, of which the natural frequency is
approximately one and a half times greater than that of the first
vibration harmonic, and of which the deformation has a nodal line
at the half-height of the blade 20, 30. Such vibrations limit the
mechanical performance of the fan 2 and of the low-pressure
compressor 3, accelerate the wear of the turbomachine and reduce
its lifetime.
[0053] As can be seen in FIG. 3a, the tangential movement by
flutter of the fan 2 blade 20 is different from that of the ferrule
32 of the low-pressure compressor 3. Indeed, the length of the
blade 20 of the fan 3 being greater than that of the low-pressure
compressor 3 blade 30, the tangential flexural moment caused by the
pulses of a fan 2 blade 20 is much greater than that caused by the
pulses of a low-pressure compressor 3 blade 30. In addition, the
stiffness of mounting within the fan 2 is different from that of
mounting within the compressor 3. With reference to FIG. 3b, this
deviation in tangential pulses is particularly visible at the
interface between the platform 25 of a fan 2 blade 20, and the
ferrule 32 knife edge seals 36.
[0054] In a first embodiment illustrated in FIG. 1, the damping
device 4 is accommodated under the platform 25 of a fan 2 blade 20.
All or a part of the fan 2 blades 20 can be equipped with such a
damping device 4, depending on the damping desired, but also the
acceptable characteristic maintenance periods.
[0055] The plurality of first radial external surfaces 40, 42, 44
is accommodated at the upper, or external portion, with respect to
the turbomachine longitudinal axis X-X, of the damping device 4.
This plurality of first radial external surfaces 40, 42, 44 is
supported with friction against the fan 2 at a plurality of
respective internal surfaces 250, 252, 254 of the platform 25 of
the fan 2 blade 20. As can be seen in FIG. 1, this plurality of
internal surfaces 250, 252, 254 delimits platform 25 bosses 251,
253 which protrude below the platform 25 in the direction of the
longitudinal axis X-X. Advantageously, the external portion of the
damping device 4 thus fits itself to the major portion of the
internal surface of the platform 25, said surface being defined by
the plurality of internal surfaces 250, 252, 254 and by the
internal surface of the bosses 251, 253. Even more advantageously,
as illustrated in FIG. 1, the damping device comprises three first
support surfaces 250, 252, 254 and the platform comprises two
bosses 251, 253. This is however not limiting, because such a
damping device 4 can be implemented under any type of blade 20
platform 25.
[0056] The second face 41 is also external to the damping device 4,
and supported with friction against the circumferential extension
34 of the ferrule 32. This ensures tangential coupling with high
stiffness between the fan 2 and the low-pressure compressor 3, so
as to reduce the tangential vibrations previously described. The
coupling is in fact the greater as the zone in which the damping
device 4 is disposed has the higher relative tangential movements
for the zero-dephasing mode considered, as illustrated in FIGS. 3a
and 3b. Typically, these relative displacements are on the order of
a few millimeters. Furthermore, the damping device 4 also
advantageously retains effectiveness on vibrational mode of the fan
2 blades 20 with non-zero dephasing.
[0057] As visible in the figures, the "stepped" structure of the
damping device 4 results from the succession, from upstream to
downstream, of the plurality of the first radial external surfaces
40, 42, 44 and from the second radial external surface 41, the
first blade 20 platform 25 being inclined with respect to the
longitudinal axis X-X. In addition, in the stepped structure of the
damping device 4, the first 40 of the first radial external
surfaces is closer to the longitudinal axis X-X than the second 42
of the first radial external surfaces, said second 42 of the first
radial external surfaces being itself closer to the longitudinal
axis X-X than the third 44 of the first radial external surfaces.
In other words, in the stepped structure of the damping device 4,
the first radial external surfaces 40, 42, 44 are each radially
more and more distant from the longitudinal axis X-X. As can be
seen in FIG. 3a, the inclination of the platform 25 advantageously
allows guiding the flow of air 5 toward the inlet of the
low-pressure compressor 3, of which the second blade 30 roots 39
are more distant from the longitudinal axis X-X than the first
blade 20 roots 29.
[0058] In a second embodiment, with reference to FIGS. 1, 4b and
5b, the damping device 4 comprises an abutment surface 46 supported
against the connection 22 between the fan 2 and the low-pressure
compressor 3, so as to accomplish the axis retention of the damping
device 4.
[0059] Advantageously as can be seen in the figures, this surface
46 forms an interior corner of the damping device 4, of which the
two edges are mutually perpendicular so as to fit it to the shape
of an attachment corner 22, such as an edge of a bolted connection
22.
[0060] In a third embodiment illustrated in FIGS. 4a, 4b, 5a and
5b, each first radial external surface 40, 42, 44, the second
radial external surface 41, and the abutment surface 46 of the
damping device 4 are formed respectively from a sacrificial plate
43, 45, 47, 49 configured to guarantee the support of said radial
external surfaces 40, 41, 42, 44, 46 of the damping device 4
against the fan 2, the low-pressure compressor 3 and the connection
22 between the fan 2 and the low-pressure compressor 3. Indeed, the
mechanical forces during operation are such that slight tangential,
axial and radial movements of the damping device 4 should be
expected. These movements are in particular due to the tangential
pulses to be damped, but also the centrifugal loading of the
assembly 1. It is necessary that these movements do not cause wear
on the blades 20 or the ferrule 32, of which the coatings are
relatively fragile. In this regard, the sacrificial plates 43, 45,
47, 49 comprise an anti-wear material, for example of the Teflon
type, or any specific composite material known to the man skilled
in the art. In addition, the sacrificial plates 43, 45, 47, 49 can
be treated by dry lubrication, for the purpose of maintaining the
value of the friction coefficient between the damping device 4 and
the ferrule 32 and/or the blade 2 platform 25. This lubrication is
for example of the MoS2 type.
[0061] For the purpose of improving the support with friction of
the damping device 4, the sacrificial plates 43, 45, 47, 49 can
also comprise an additional coating 430, 450, 470, 490, as can be
seen in FIG. 4b. Generally, such a coating 430, 450, 470, 490 is
configured to reduce the friction and/or the wear of the engine
parts between the plate 42 and the rotor modules 2, 3. This coating
430, 450, 470, 490 is for example of the viscoelastic type. Such a
coating 430, 450, 470, 490 then advantageously comprises a material
having properties similar to those of a material such as the range
having the commercial designation "SMACTANE.RTM.," for example a
material of the "SMACTANE.RTM. 70" type. Another means of
increasing the tangential stiffness of the assembly 1 is to
sufficiently preload the viscoelastic coating 430, 450, 470, 490,
for example during the assembly of the assembly 1, so that the
relative tangential movement between the blade 20 and the ferrule
32 is transformed into viscoelastic shear of the coating 430, 450,
470, 490 alone.
[0062] Alternatively, this coating 430, 450, 470, 490 is of the
dissipative and/or viscoelastic and/or damping type. The
dissipative coating 430, 450, 470, 490 then comprises a material
chosen from those having mechanical properties similar to those of
Vespel, of Teflon or of any other material with lubricating
properties. More generally, the material has a coefficient of
friction comprised between 0.3 and 0.07.
[0063] In this manner, the damping device 4 is not too flexible
tangentially. Too high a flexibility would not allow the damping of
the mode with zero dephasing, because the relative movements of the
fan 2 and of the low-pressure compressor 3 would lead to friction
and/or oscillations between a "stuck" state and a "slipping" state
of the damping device 4.
[0064] These additional coatings 430, 450, 470, 490 are applied by
gluing to the sacrificial plates 43, 45, 47, 49.
[0065] Advantageously, as can be seen in FIG. 5b, the plate 43
forming the farthest downstream radial external surface 40 has a
variable thickness along the turbomachine longitudinal axis X-X.
Preferably, the farthest upstream plate 43 portion is thicker than
the plate 43 portion farthest downstream. This allows optimizing
the distribution of forces in the coupling between the fan 2 and
the low-pressure compressor 3.
[0066] In a fourth embodiment illustrated in FIG. 4a, damping by
tangential coupling can be adjusted by controlling the mass of the
damping device 4, which influences the shear inertia. This control
involves modifications of the mass of the damping device 4. This
mass can be modified in all or a part of the damping device 4,
typically by making bores 7 to lighten it, and/or adding one or
more inserts 8, metallic for example, to weigh it down.
[0067] Advantageously, the combination of the third and fourth
embodiment allows adjusting the contact forces between the damping
device 4 and the fan 2 and the low-pressure compressor 3. Indeed,
contact forces that are too high between the fan 2 blade 20 and the
damping device 4 would limit the dissipation of vibrations during
operation.
[0068] In a fifth embodiment illustrated in FIGS. 5a and 5b, the
farthest downstream first radial external surface 40 among the
plurality of first radial external surfaces 40, 42, 44, and the
second radial external surface 41, are dug by a bore 400 so as to
form a U-shaped radial external surface 40. The bore 400 can pass
through all or a part of the downstream portion of the damping
device 4.
[0069] This bore 400 allows an increase in the flexibility of the
downstream portion of the damping device 4. In addition, the U
configuration allows adapting the support with friction of the
radial external surface 40 to the deviations between two
circumferentially successive blades 20. Thus, the static redundancy
of the damping device 4 is advantageously reduced.
[0070] Different embodiments of the assembly 1 according to the
invention have been described in the case where the first rotor
module 2 is a fan, and the second rotor module 3 is a low-pressure
compressor.
[0071] This, however, is not limiting, because the first rotor
module 2 can also be a first, high- or low-pressure, compressor
stage, and the second rotor module 3 a second stage of said
compressor, successive to the first compressor stage, upstream or
downstream of the latter. Alternatively, the first rotor module 2
is a first, high- or low-pressure, turbine stage and the second
rotor module 3 a second stage of said turbine, successive to the
first turbine stage, upstream or downstream of the latter.
* * * * *